The present invention relates to a turbine engine. In particular, the invention relates to a combustor seal system for a gas turbine engine.
A turbine engine ignites compressed air and fuel in a combustion chamber, or combustor, to create a flow of hot combustion gases to drive multiple stages of turbine blades. The turbine blades extract energy from the flow of hot combustion gases to drive a rotor. The turbine rotor drives a fan to provide thrust and drives a compressor to provide a flow of compressed air. Vanes ahead of, and interspersed between, the multiple stages of turbine blades align the flow of hot combustion gases for an efficient attack angle on the turbine blades.
In most instances, a portion of the flow of compressed air flows around the combustor to cool the combustor and to cool components downstream from the combustor. Leakage of the compressed air into the flow of combustion gases may deprive the downstream components of the cooling necessary to prolong their service life. The loss of compressed air into the flow of combustion gases may also lead to increased compressed air requirements. Energy expended on compressing air that leaks into the flow of combustion gases is not available to produce thrust and engine efficiency is reduced.
Seals are generally employed between components to prevent leakage of the compressed air into the flow of combustion gases. Improvements in the effectiveness of such seals can reduce leakage of compressed air into the flow of combustion gases, and improve the overall efficiency of the turbine engine.